The present invention concerns gas-turbine-engine rotors and, more particularly, axial-locking means for the blades mounted in axial openings on the disk periphery.
When the rotor-blades of gas turbine engines are fastened in axial openings--that is, openings extending parallel to the axis of the gas-turbine engine, or slightly apart from a parallel to the axis of the gas turbine engine, two assembly problems do arise.
A first problem is to achieve the simplest possible locking of each blade, but also in the most reliable manner, and in such a way that disassembly of a blade or of the entire set may also be simplified.
The second problem concerns the hermeticity between the upstream and downstream sides of a disk and is crucial for the gas-turbine-engine compressors. If the spaces between the alveolar bottoms of the disk and the blade roots are not suitably masked, substantial downstream air volume may pass under the blade roots and recirculate to the compressor upstream side, whereby its compression ratio shall be lowered and the overall efficiency of the gas-turbine engine shall be prohibitively degraded.
French patent document A 2,603,333 in the name of applicant describes a gas-turbine-engine rotor, in particular for aviation purposes, which comprises at least one disk bearing a set of blades of which the roots are mounted in broached alveoli along the disk periphery and along an axis parallel with or slightly inclining to a parallel to the longitudinal engine axis. The blade roots are equipped with a means for wedging the blades onto the upstream disk side and with a rear lip fitted with a transverse groove radially pointing to the disk axis, the disk itself comprising a circular groove radially pointing to its periphery. The rotor includes a means for locking the blades axially downstream on the disk while simultaneously the hermeticity between said blade roots and the disk's alveolar bottoms is assured, said means for locking the blades downstream onto the disk consisting of two split rings of which the first at least cooperates simultaneously with the grooves of the blade lips and the circular disk groove, the sum of the thicknesses of the first and second rings being equal to the thickness of the grooves of the blade lips.
This document represents the state of the art transcended by the present invention.
In this prior art, the blades required a groove deep and wide enough to receive the two rings.
As regards gas turbine engines presently the object of research, illustratively turbojet engines with rapid propellers for which the compressor rotors evince very small diameters, the height underneath the platforms must be minimized and mass gains must be realized with respect to the compressor blades, and the above solution is inapplicable because too bulky.